Aircraft and missile afterbody flow control device and method of controlling flow

ABSTRACT

An afterbody flow control system is used for aircraft or missile flow control to provide enhanced maneuverability and stabilization. A method of operating the flow control system is also described. The missile or aircraft comprises an afterbody and a forebody; at least one activatable flow effector on the missile or aircraft afterbody; at least one sensor having a signal, the at least one sensor being positioned to detect forces or flow conditions on the missile or aircraft afterbody; and a closed loop control system; wherein the closed loop control system is used for activating and deactivating the at least one activatable flow effector based on at least in part the signal of the at least one sensor.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.11/415,534, filed on May 2, 2006 and issued Jun. 5, 2012 as U.S. Pat.No. 8,191,833 B1, which is a continuation of patent application Ser. No.10/750,422 filed Dec. 20, 2003 and issued as U.S. Pat. No. 7,070,144 onJul. 4, 2006, which is a continuation-in-part of U.S. patent applicationSer. No. 10/336,117 filed Jan. 3, 2003 and issued as U.S. Pat. No.6,685,143 on Feb. 3, 2004.

The U.S. Government has a paid-up license in this invention and theright in limited circumstances to require the patent owner to licenseothers on reasonable terms provided for by the terms of grantDAAE30-02-C-1052 awarded by the U.S. Army.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to an aerodynamic flow control system andmore particularly to aircraft or missile afterbody flow control systemfor enhanced maneuverability and stabilization at low angles of attack.The present invention further relates to a method of operating the flowcontrol system.

2. Technical Background

Traditional aircraft and missile maneuvering technologies utilizeconventional control surfaces such as tail-fins and canards to providecontrol and stability through all phases of an aircraft or missile'sflight path. These control surfaces require a significant payload andvolume to house the control actuation system for these conventionalcontrol surfaces, which includes heavy servomotors, thereby imposingsignificant limitations on the aircraft or missile aerodynamicperformance. In addition, these hinged-control surfaces also reduce theeffective payload, maximum achievable range, and lethality of missilesand aircraft.

Conventional missile and aircraft control techniques are not capable ofmeeting new multi-mission highly accurate, long-range fire requirementsthat are needed to ensure the multi-target engagement capabilities ofmissiles or aircraft, particularly “smart” missiles and aircraft. Inaddition, with missiles the overall cost of the control system per roundneeds to be minimum without affecting the aerodynamic efficiency of themissiles considering their subsistence period once deployed. The mostpopular choice of steering control for missiles or aircraft is afterbodyor tail-based control due to its flexibility in modifying aerodynamicdesigns without affecting other surfaces (canards, wings, etc.) as theyfall out of its zone of influence. The major disadvantages of afterbodyor tail-based control is space restriction, i.e., control surfaces mustbe located in an annular space around the throat of the propulsionnozzle, increased weight, and drag from exposed surfaces. Theconventional control surfaces necessitate hinges, which increase theoverall weight-induced aerodynamic drag, as well as the complexity ofthe propulsion system.

In view of the foregoing disadvantages of presently available controlsurfaces, it is desirable to develop a missile or aircraft aerodynamicafterbody control system flow control system that provides the necessaryforces for missile or aircraft control with limited or no use of hingedcontrol surfaces. It is further desirable to develop a missile oraircraft with an aerodynamic control system for maneuvering at lowangles of attack. It is still further desirable to develop a missile oraircraft aerodynamic flow control system that is highly compact andlightweight with the ability of being deactivated when not required.

SUMMARY OF THE INVENTION

The present invention relates to an aerodynamic flow control system andmore particularly to aircraft or missile with an afterbody flow controlsystem for enhanced maneuverability and stabilization. The presentinvention further relates to a method of operating a missile or aircraftwith such an aerodynamic flow control system.

In one embodiment, the present invention includes a missile or aircraftcomprising an afterbody and a forebody; at least one activatable floweffector or active flow control device on the missile or aircraftafterbody; at least one sensor each having a signal, the at least onesensor being positioned to detect a force or flow separation on themissile or aircraft afterbody; and a closed loop control system; whereinthe closed loop control system is used for activating and deactivatingthe at least one activatable flow effector or active flow control devicebased on at least in part the signal of the at least one sensor.

In another embodiment, the present invention includes a flow controlsystem for a missile or aircraft afterbody comprising at least oneactivatable flow effector or active flow control device; an inertialmeasurement unit having an output; and a closed loop control system;wherein the closed loop control system is used for activating anddeactivating the at least one activatable flow effector or active flowcontrol device based at least in part on the signal of the output of theinertial measurement unit.

In still another embodiment, the present invention includes a method ofmaneuvering a missile or aircraft afterbody comprising the steps ofactivating at least one activatable flow effector or active flow controldevice to create side forces on the missile or aircraft afterbody;estimating or determining side forces or flow separation on a missile oran aircraft afterbody based at least in part on a signal from at leastone sensor, the at least one sensor being positioned to detect forces orflow separation on the missile or aircraft afterbody; and deactivatingthe at least one activatable flow effector or active flow control devicein response to changed forces or a flow condition.

In still another embodiment, the present invention includes a missile oraircraft having a yawing moment comprising at least one activatable floweffector or active flow control device on the missile or aircraftboattail or tail fins wherein the at least one activatable flow effectoror active flow control device is used to change the yawing moment of themissile or aircraft.

In still another embodiment, the present invention includes a missile oraircraft having a pitching moment comprising at least one activatableflow effector or active flow control device on the missile or aircraftboattail or tail fins wherein the at least one activatable flow effectoror active flow control device is used to change the pitching moment ofthe missile or aircraft.

In still another embodiment, the present invention includes a missile oraircraft having rolling moment comprising at least one activatable floweffector or active flow control device on the missile or aircraftboattail or tail fins wherein the at least one activatable flow effectoror active flow control device is used to change the rolling moment ofthe missile or aircraft.

In still another embodiment, the present invention includes a missile oraircraft comprising a missile or aircraft having a boattail, theaircraft having drag and the boattail having at least one activatableflow effector or active flow control device wherein the at least oneactivatable flow effector or active flow control device is used toreduce the drag of the aircraft.

Additional features and advantages of the invention will be set forth inthe detailed description which follows, and in part will be readilyapparent to those skilled in the art from that description or recognizedby practicing the invention as described herein, including the detaileddescription which follows, the claims, as well as the appended drawings.

It is to be understood that both the foregoing general description andthe following detailed description are merely exemplary of theinvention, and are intended to provide an overview or framework forunderstanding the nature and character of the invention as it isclaimed. The accompanying drawings are included to provide a furtherunderstanding of the invention, and are incorporated in and constitute apart of this specification. The drawings illustrate various embodimentsof the invention, and together with the description serve to explain theprinciples and operation of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1. Schematic view of one embodiment of a missile having a forebodyand afterbody with activatable flow effectors or active flow controldevices, and sensors mounted on the tail fins and boattail.

FIG. 2. Schematic view of one embodiment of an aircraft having aforebody and afterbody with activatable flow effectors or active flowcontrol devices, and sensors mounted on the tail fins and boattail.

FIG. 3. a) Perspective view of one embodiment of the afterbody sectionof a missile or aircraft having activatable flow effectors or activeflow control devices, and sensors mounted therein; b) Sectional view oftail fin along plane A-A′ shown in FIG. 3a ).

FIG. 4. Perspective view of one embodiment of a module containing aco-located sensor, and a) a deployable flow effector (deployed) and b) adeployable flow effector (retracted).

FIG. 5. Sectional view of one embodiment of the afterbody section of amissile or aircraft a) fin and b) boattail.

FIG. 6. Sectional view of one embodiment of a deployable flow effector.

FIG. 7. Sectional view of deployable flow effector shapes.

FIG. 8. Sectional view of another embodiment of a deployable floweffector.

FIG. 9. Depiction of an embodiment using a plasma actuator.

DESCRIPTION OF THE PREFERRED EMBODIMENT

The present invention relates to an aerodynamic flow control system andmore particularly to aircraft or missile with an afterbody flow controlsystem for enhanced maneuverability and stabilization. The afterbody ofthe missile or aircraft for the present invention is defined as the backhalf of the length of the missile or aircraft, and more particularly asthat section of fuselage behind the wings including but not limited tothe tail section, i.e., the boattail and tail fins. The boattail is therear portion of a missile or aircraft having decreasing cross-sectionalarea towards the rear. Preferably, the afterbody is the back 25% of thelength of the missile or aircraft, and most preferably the afterbody isthe tail section of the missile or aircraft.

The present invention most preferably relies on a successful marriage ofthe boattail geometry with the activatable flow effectors or active flowcontrol devices incorporated within the geometry for aerodynamiccontrol. The geometry of the boattail plays a critical part in how wellthe activatable flow effectors or active flow control devices perform.The onset of boattail and the taper angle is preferably optimized forthe desired speed application so as the flow condition in the vicinityof the boattail is amenable to flow control devices. The boattail taperdesign, which depends on the flow conditions as well as fluid parameterssuch as the viscosity, density, temperature and pressure, is preferablyoptimized for the given application.

Additionally or alternatively, the same holds true for the tail-finoptimization as well. In other words, the present invention mostpreferably relies on a successful marriage of the tail-fin geometry withthe flow control devices incorporated within the geometry foraerodynamic control. The geometry of the tail-fins plays a critical partin how well the flow control devices perform. The design of the tail-finis preferably optimized for the desired speed application.

The activatable flow effector or active flow control devices of thepresent invention are electromechanical devices that can be used tocreate disturbances in the flow over the surface of the missile oraircraft. Preferably, the activatable flow effector or active flowcontrol devices induce small disturbances or perturbances in thevicinity or close proximity to the activatable flow effector or activeflow control device. Further preferably, the activatable flow effectoror active flow control device is flush or nearly flush, whendeactivated, with the surface of the missile or aircraft to which it hasbeen installed thereby creating little or no drag on the missile oraircraft. Still further preferably, the activatable flow effector oractive flow control devices have no hinged parts or surfaces. Theactivatable flow effector or active flow control devices of the presentinvention include but are not limited to active vortex generators, whichare deployable including but not limited to flow deflectors, balloons,microbubbles, and dimples, or create active pressure active regions bysuction or air pressure; synthetic jets including zero-net-masssynthetic jets; pulsed vortex generators; directed jets; vortexgenerating devices (fluidic and mechanical); plasma actuators, includingweakly ionized plasma actuators; wall turbulators; porosity includingbut not limited to reconfigurable, inactive and active; microactuators;and thermal actuators. The present invention further relates to a methodof operating the flow control system.

For stabilization and maneuverability of the missile or aircraftafterbody, the flow control system relies on the effectiveness of theactivatable flow effector or active flow control devices in generatingon-demand forces or flow conditions at different regions around themissile or aircraft afterbody to create the desired flow effectincluding but not limited to stabilization or maneuverability of theaircraft or missile. Flow condition may be defined as a mode or a stateof a fluid with a characteristic behavior. These states may be definedby physical factors such as: density, velocity, temperature, viscosity,and pressure. Flow conditions can be of several types. Flows may besteady (not time-varying) or unsteady (time-varying), one-dimensional,two-dimensional or three-dimensional. These conditions are different forcompressible and incompressible flows. Other conditions include, but arenot limited to, laminar, transitional, turbulent, attached, reattached,and shear flow.

The flow control system for the missile or aircraft afterbody can beused at both low and high angles of attack. The activatable floweffector or active flow control devices of the present invention areactive micro-vortex generators that effectively control the pressuredistribution along the afterbody of the missile or aircraft, yieldingforces and yawing, rolling, and pitching moments for controlling of yaw,roll or pitch on the missile or aircraft body. The activatable floweffector or active flow control devices of the present inventionpreferably are deployable flow effectors or other types of micro-vortexgenerators. Activatable flow effectors or active flow control devices ofthe present invention are flow effectors that are activated to generatefluid flow disturbances in the vicinity of the flow effector, and thatcan be deactivated when not needed. Preferably, the activatable floweffector or active flow control devices of the present invention can beoperated at high frequencies. Further preferably, the activatable floweffector or active flow control devices are capable of being cycled atfrequencies of at least about 1 Hz, more preferably at frequencies of atleast about 20 Hz, even more preferably at frequencies of at least about60 Hz, even more preferably at frequencies of at least about 100 Hz, andmost preferably at frequencies of at least about 250 Hz. One type ofactive flow control device or activatable flow effector is a deployableflow effector, which is described in more detail in the variousembodiments in the Figures below. The frequencies at which the activeflow control device or activatable flow effector of the presentinvention are cycled may be determined based in part on a number offactors including but not limited to the desired flow effect, autopilotfrequency response characteristics, missile or aircraft dynamics, andmissile or aircraft environmental conditions.

Some of the other types of activatable flow effectors or active flowcontrol devices not shown in the Figures (but described in more detailin U.S. Pat. No. 6,302,360 B1 to Ng which is herein incorporated byreference) include but are not limited to spaced apart valves that arepositioned at inlets of a vacuum or pressure chamber, or are connectedby pneumatics to a vacuum or pressure source. Preferably, the valvescontain a flap that operates to open and close the valves as directed byelectrostatic forces. Other valve configurations can also be used. Whenthe valves are opened, the negative pressure from the vacuum chamber orsource causes withdrawal of air from the surface of the missile oraircraft forebody through the surface orifices. Therefore, it can beseen that the opening of the valves causes the pressure active region togenerate a net inflow of air, resulting in the generation of vortices,that can be used to control the airflow around the afterbody surface ofthe missile or aircraft for desired control effect such as improvedmaneuverability and/or stability. Similarly, when the valves are open toa positive pressure chamber or source, a net outflow of air is generatedresulting in the generation of vortices, which also act beneficially toreattach the air flow to the afterbody surface of the missile oraircraft. For purposes of this invention activatable flow effectors oractive flow control devices include any type of device or article knownto those skilled in the art or discovered at a later point that is usedto assist in the reattachment of airflow to a missile or aircraft'ssurface. Preferably, the activatable flow effectors or active flowcontrol devices of the present invention are deployable flow effectors.Further preferably, the missile or aircraft of the present invention hasat least about 4 activatable flow effector or active flow controldevices, more preferably at least about 6 activatable flow effector oractive flow control devices, even more preferably at least about 8activatable flow effector or active flow control devices, still evenmore preferably at least about 50, and most preferably at least about200. The applicants further incorporate by reference U.S. patentapplication Ser. Nos. 10/336,114, and 10/336,113.

Referring now to FIG. 1, there is shown a schematic view of oneembodiment of a missile 10 having a forebody 18 and an afterbody 13. Theafterbody 13 has at least one activatable flow effector or active flowcontrol device 12. The afterbody further has at least one sensor 14. Thesensor is positioned to detect flow separation from the flow surface 16on the missile 10 afterbody 13. In this embodiment, activatable floweffector or active flow control devices 12 as well as sensors 14 areincorporated into the boattail 17 and tail fins 19 of the missile. Theafterbody 13 of this specific embodiment has a number of flow effectors12 and sensors 14 mounted in the afterbody 13 (or tail section) therein.Furthermore in this specific embodiment, the individual flow effectors12 and individual sensors 14 are in close proximity with respect to eachother. The fluid boundary layer is a thin layer of viscous flowexhibiting certain pressure variation characteristics and fluid dynamicsthat affect the operation of the flow surface 16. The fluid is generallyair. The flow surface 16 for purposes of the present invention is theafterbody of a missile or an aircraft. The missile 10 shown has midbodyfins 6.

FIG. 2 is a schematic view of one embodiment of an aircraft 20 adaptedwith the vortex generating system 22 of the present invention. Theairplane can be any type of aircraft, including commercial, military andspace vehicles. The aircraft 20 includes a fuselage 21, a tail 23, wings24, forebody (nose) 18, and afterbody 13. In this specific embodiment,the individual activatable flow effectors or active flow control devices12 and individual sensors 14 are also mounted in close proximity withrespect to each other on the afterbody 13 of the aircraft 20 includingon the aircraft's 20 tail fins 19 and boattail 17.

The sensor(s) of the present invention include but are not limited to adynamic pressure sensor, shear stress sensor (hot film anemometer, adirect measurement floating-element shear stress sensor), inertialmeasurement unit or system, and other sensors known to those skilled inthe art whose signal could be used to estimate or determine flowcondition such as separation on the surface of the missile or aircraft,which would function as a trigger point for actuating the activatableflow effectors or active flow control devices. The sensors of thepresent invention are used to determine or estimate flow separation. Aninertial measurement unit, for example, is a sensor that would notdirectly measure forces or flow separation, but could be used toestimate or predict separation. The preferred sensor of the presentinvention is a pressure sensor. The pressure sensor is used to predictor sense flow separation. The pressure sensor can be any type of sensorsuitable for measuring the pressure at the flow surface. The pressuresensor can be, for example, a piezoelectric device that generates anelectric signal in response to a sensed pressure, a shape memory alloydevice, or any other pressure sensor or transducer known to thoseskilled in the art. Preferably, the ratio of flow effectors to sensor isless than about 100:1, more preferably less than or equal to about 50:1,still preferably less than or equal to about 20:1, even more preferablyless than or equal to about 3:1, still even more preferably less than orequal to about 2:1, and most preferably less than or equal to 1:1. Thehigher the concentration of pressure sensors to flow effectors, the moreredundancy can be built into the system utilizing the present invention.Most preferably the sensor is a flush, surface-mounted, diaphragm-typepressure sensor. The at least one sensor 14 has a signal that is used atleast in part by a controller (not shown) to activate and deactivate theat least one activatable flow effector or active flow control device 12.

In addition to pressure sensors, various embodiments of the presentinvention may also include a means for determining the relative spatialorientation of the flow effectors and/or sensors with respect the flowseparation on the missile or aircraft body. This means would includeutilizing the output of an inertial measurement unit and other systems,which could be used to determine the missile or aircraft orientation. Aninertial measurement unit provides six-degree-of-freedom motion sensingfor applications such as navigation and control systems. Angular rateand acceleration are measured about three orthogonal axes.

FIG. 3a ) is a perspective view of one embodiment of the afterbodysection and more particularly the tail section 23 of a missile oraircraft having an activatable flow effector or active flow controldevice 12 and sensors 14 mounted therein. The missile or aircraftafterbody of the present invention can be designed with activatable floweffector or active flow control devices 12 and sensors 14 mounted in thetail fins 19 and/or the boattail 17 to provide for improved stability,maneuverability, or controllability with the present flow controlsystem. This flow control system is designed to provide for a variety ofmoments about the aircraft, which allow for both flow separation andflow attachment and result in improved stability and maneuverability.These moments can be used to change the drag, the yaw, the lift, theroll, the pitch and the thrust of the missile or aircraft. FIG. 3b ) isa sectional view of section A-A′ of a missile or aircraft tail fin 19 asshown in FIG. 3a ). FIG. 3b ) shows the fluid flow 28 around a missileor aircraft tail fin 19 at a section A-A′ in the proximity of theactivatable flow effector or active flow control devices 12, and thevarious embodiments of the resultant flow separation and reattachmentwith activation and deactivation of the activatable flow effector oractive flow control device 12 of the present invention.

In FIG. 4, there is shown a perspective view of one embodiment of amodule containing a co-located sensor, and a) an activatable, deployableflow effector (deployed) and b) an activatable, deployable flow effector(refracted). In this particular embodiment, the module 32 contains anactivatable, deployable flow effector 12 and a pressure sensor 14. Theactivatable, deployable flow effector 12 being capable of being deployedinto and retracted from, respectively, the fluid boundary layer flowingover the flow surface of the missile or aircraft forebody wherein themodule 32 is employed. The deploying and retracting can be accomplishedusing any device such as pneumatic pressure, hydraulic pressure, vacuum,a mechanical device such as a solenoid valve, a microelectromechanicaldevice, any combination thereof or the like. The module 32 may or maynot include a controller (not shown) internal to the module. Thepressure sensor 14 is connected to the controller (not shown). If thecontroller (not shown) is not internal to the module 32 then the module32 preferably further comprises a link between pressure sensor 14 andthe controller, and another link between the controller (not shown) anddeploying means (not shown). The controller (not shown) is programmed tooperate the deploying and retracting means in response to specificpressure conditions sensed at the flow surface (not shown). Thecontroller (not shown) can be any device, such as a computer, suitablefor gathering information from the pressure sensors 14 and directing theactivation of the activatable flow effector or active flow controldevices 12. Where a number of activatable flow effectors or active flowcontrol devices 12 and/or pressure sensors 14 (or modules 32) areemployed, the controller (or controllers) (not shown) can be programmedand connected to integrate each of the activatable flow effectors oractive flow control devices 12, pressure sensors 14, and modules 32 sothat the output from all of the regions will be coordinated to enhanceand possibly optimize the stabilization and maneuverability of a missileor an aircraft forebody. Specific patterns of deployment and/orretraction of the flow effectors 12 can be determined to handle avariety of routine events and also incorporated into the control scheme.

FIG. 5 is a sectional view of one embodiment of the afterbody section(tail section) of a missile or aircraft having an activatable floweffector or active flow control device 12 and sensor 14 mounted therein.In FIG. 5, the two activatable flow effectors or active flow controldevices 12 shown in this cross-section are movably attached by anattachment means, i.e., a hinge 91, to a base structure 82. Theactivatable flow effectors or active flow control devices 12 aredeployable flow effectors. The activatable flow effectors or active flowcontrol devices 12 are further movably attached to a piston 84. Thepiston 84 moves within a cylinder 86 in response to a pressure source(not shown) applied via a pneumatic system 89 against an elastomericsheet 81 to move the pistons 84 and in return to deploy and retract theflow effectors 12. The piston 84 also is connected to a biasing means87, i.e., a spring, to return the piston 84 to its original positionupon removing the pressure source, and therefore retracting thesedeployable flow effectors 12. In this particular embodiment, thepressure is applied to the piston 84 via a pressure inlet/outlet 88.Also shown in this particular embodiment are seals in the form ofO-rings 83 to seal the pneumatic system (not shown) of the pressuresource (not shown), and two sensors 14. The sensors 14 are connected vialeads to a controller (not shown). The pressure source (not shown) isalso connected to the controller (not shown).

FIG. 6 is a sectional, detailed view of a module 32 (as shown in FIG. 4)with an activatable deployed flow effector 12. In FIG. 6, theactivatable deployed flow effector 12 is movably attached to the upperportion 48 of the housing 46 of the module 32 and is attached to thelower portion 50 of the housing 46 of the module 32 by at least twofasteners 40. The upper portion 48 of the housing 46 mates with thelower portion 50 with a sealing ring (not shown) and a sealable,flexible element 44 there between. The flow effector 12 is deployed bypressure being applied to the flexible element 44. The flow effector 12has a biasing means (a spring) 47 which attaches at one end to the upperportion 48 of the housing 46 and at the other end to the base 54 of floweffector 12. Directly beneath the flow effector 12 is a valve 43, whichopens and closes to allow for the application of fluid or gas pressurefrom a pressure source (not shown) to be applied to the flexible element44 through a pneumatic pathway 52. A pressure sensor 14 senses fluidflow at or near the surface over which the fluid is flowing. Preferablythe pressure sensor is at the surface of the airfoil, and mostpreferably it is flush with such surface. The pressure sensor 14 can beany pressure sensor, but advantageously is a microelectromechanical(MEMS) based or piezoelectric based sensor. MEMS devices are smallmechanical/electrical systems that perform small-scale tasks thatprovide large-scale effects. MEMS devices are generally manufacturedusing batch microfabrication technology, the same manufacturingtechnology used to make integrated circuits (IC's). Consequently, manyof the same benefits of IC manufacturing are applicable to MEMSmanufacturing including high reliability, high yield, and low cost.Furthermore, since IC's and MEMS are both silicon-based technologies andare fabricated using similar techniques, it is relatively easy to mergemicroelectronics and micromechanical elements onto the same substrates.Electrostatic actuated MEMS devices have two dominating advantages ascompared to other actuation mechanisms, which are high bandwidth and lowpower consumption. The sensor transmits a signal, in this case avoltage, but it is understood to one skilled in the art that the signalcan be other than voltage, including, but not limited to, current,pressure, hydraulic, or optical. The signal corresponds to the pressureit senses.

The pressure sensors 14 (or other sensors) are connected to a controller42 internal to the module 12 (or optionally external to the module). Thecontroller 42 can be, for example, a proportional-integral-derivative(PID) controller, an adaptive predictive controller, or an adaptivepredictive feedback controller. The controller of the present inventionis preferably a closed loop control system. The controller can be usedto minimize side forces or to create commanded side forces on themissile or aircraft forebody. The pressure sensor transmits a signal tothe controller 42 through the electrical connection 38 (in practicalapplication, multiple pressure sensors 14 send multiple signals to thecontroller 42). The controller 42 processes the signals to determine,through mathematical modeling, the dynamics of the flow surface. Suchdynamics include boundary layer separation and stall. It is thepredictive ability of the controller 42 that provides for this functionand expands this system from being merely responsive. This is especiallyadvantageous for dynamic systems, which are nonlinear and time varyingand operating in challenging environments. The controller 42 produces anoutput signal to a monitor, recorder, alarm and/or any peripheral devicefor alarming, monitoring, or in some manner, affecting or precluding thedynamics upon its incipience. Advantageously, the controller 42 is theORICA controller, an extended horizon, adaptive, predictive controller,produced by Orbital Research Inc. and patented under U.S. Pat. No.5,424,942, which is incorporated herein by reference. Under certainconditions, the controller 42 (or optionally an external controller),which is connected via electrical connection 51 to the valve 43, causesthe valve 43 to open, thereby resulting in the deployment of the floweffector(s) 12.

The closed loop control system of the present invention not onlyreceives input in part from the sensors, but also can be set up toreceive input from a number of other sources. These sources can includebut are not limited to the autopilot, crash avoidance, or steeringsystems on an aircraft, or similar systems or non-integral, non-internalcommand control systems used to re-program a missile in flight. Themissile or aircraft can be maneuvered or stabilized using the flowcontrol system based in part on the sensors input and in part (ifnecessary) on new input from, for example, the autopilot into the closedloop control system to activate or deactivate the flow effectors asrequired.

Preferably, the pressure source (or other deployment and/or retractionmeans) is internal to the module 12. The sealable, flexible element 44referred to above can be made of a single polymer or a combination ofpolymers. The pressure source can be air bled from an aircraft turbineengine, a pressurized gas cartridge, or pressurized fluid. The biasingmeans is employed to urge the sealable, flexible element 44 towards itsquiescent state after pressure is removed or reduced. The biasing meanscan be any device or spring-like means, such as a vacuum or pressuredevice, a mechanical device, or an electromechanical device.

The deployable portion of the activatable, deployable flow effectorsshown in the previous Figures are small mechanical tabs preferably madefrom epoxy glass-fabric, and deactivate to assume a position underneaththe skin surface of the missile or aircraft in their retracted state.Several examples of various embodiments of the flow effectors are shownin FIG. 7. a, b, c, d, e, f, and g. These cross-sectional viewsdemonstrate that rectangular 72, triangular 74, irregular 76,semi-circular 78, a fence 73, a separated (or picket) fence 75 and ahalf-cylinder 77. The present invention is, however, not limited tothese shapes and it is envisioned that any shape of activatable,deployable flow effector known presently or conceived of in the futureby those skilled in the art may be used. Other types of deployable floweffectors that can be used include but are not limited to bumps,dimples, and tubes. Upon controlled activation, the activatable floweffectors or active flow control devices (deployable or other)manipulate the forebody of the missile or aircraft's vortical flow fieldto generate the desired forces or flow separation. Single flow effectorsor combinations of flow effectors can be activated either statically orcycled at a varying frequency (oscillated) to obtain a desired sideforce or yawing moment. Varying frequency or oscillation of the floweffectors includes but is not limited to pulse width modulation or othertechniques known to those skilled in the art.

FIG. 8 is a sectional view of another embodiment of an activatable,deployable flow effector. In FIG. 7, the activatable flow effector oractive flow control device 12 is a deployable flow effector. Theactivatable, deployable flow effector 12 is further movably attached toa camshaft 94. The camshaft 94 moves in response to an electric motor 96to deploy and retract the flow effector 12. The motor is connected to acontroller 42. The controller 42 activates and deactivates thedeployable flow effector in response to at least in part the signal fromthe sensor 14.

FIG. 9 shows missile or aircraft 801 having an afterbody 802 and aforebody 803, afterbody 802 having at least one plasma actuator 804. Theat least one plasma actuator 804 may be used to reduce the drag ofaircraft or missile 801 by activating or deactivating plasma actuator804 to reattach flow to the surface of the missile or aircraft 801.Plasma actuator 804 may be used to change a pitching or yawing moment ofmissile or aircraft 801. Missile or aircraft 801 further comprises atleast one sensor 805 having a signal. As illustrated, sensor 805 ispositioned to measure, estimate or predict a force or flow condition onthe missile or aircraft's afterbody 802. As illustrated, sensor 805 isco-located with plasma actuator 804. Missile or aircraft 801 furthercomprises closed-loop control system 806, which is used for activatingand deactivating the at least one plasma actuator 804 based at least inpart on the electrical signal of the at least one sensor 805.Closed-loop control system 806 comprisesproportional-integral-derivative (PID) controller 807.

It will be apparent to those skilled in the art that variousmodifications and variations can be made to the present inventionwithout departing from the spirit and scope of the invention. Thus, itis intended that the present invention cover the modifications andvariations of this invention provided they come within the scope of theappended claims and their equivalents.

What is claimed:
 1. A missile or aircraft comprising: an afterbody aforebody and at least one sensor having an electrical signal; themissile or aircraft afterbody having a boattail and tail fins; themissile or aircraft experiencing drag during flight; and at least oneelectromechanical activatable flow effector or electromechanical activeflow control device on the boattail or tail fins, wherein the at leastone sensor is positioned to measure, estimate or predict a force or flowcondition the missile or aircraft's afterbody, the at least oneelectromechanical activatable flow effector or electromechanical activeflow control device is configured and located on the boattail or tailfins to reduce the drag of the missile or aircraft by activating anddeactivating the electromechanical activatable flow effector orelectromechanical active flow control device to reattach flow to thesurface of the missile or aircraft, and the activation or deactivationof the activatable flow effector or active flow control device iscapable of reattaching detached flow to the surface of the missile oraircraft.
 2. The missile or aircraft in claim 1, further comprising aclosed-loop control system, wherein the closed-loop control system isused for activating and deactivating the at least one activatable floweffector or active flow control device based at least in part on theelectrical signal of the at least one sensor.
 3. The missile or aircraftin claim 2, wherein the closed-loop control system activates anddeactivates the at least one activatable flow effector or active flowcontrol device to dreate command forces on the afterbody to maneuver themissile or aircraft.
 4. The missile or aircraft in claim 1, wherein theat least one activatable flow effector or active flow control device isa deployable flow effector.
 5. The missile or aircraft in claim 4,wherein the at least one activatable flow effector or active flowcontrol device, when deactivated, is flush, or nearly flush, with thesurface of the missile or aircraft.
 6. The missile or aircraft in claim1, wherein the at least one sensor is an inertial measurement unit.
 7. Amissile or aircraft comprising: an afterbody and a forebody; and atleast one plasma actuator, wherein the at least one plasma actuator isconfigured and located on the missile or aircraft afterbody to reducethe drag of the aircraft or missile by activating or deactivating theplasma actuator to reattach flow to the surface of the missile oraircraft, and wherein the activation or deactivation of the plasmaactuator is capable of reattaching detached flow to the surface of themissile or aircraft.
 8. The missile or aircraft in claim 7, furthercomprising at least one sensor having a signal, The at least one sensorbeing positioned to measure, estimate or predict a force or flowcondition on the missile or aircraft's afterbody.
 9. The missile oraircraft in claim 8, further comprising a closed-loop control system,wherein the closed-loop control system is used for activating anddeactivating the at least one plasma actuator based at least in part onthe electrical signal of the at least one sensor.
 10. The missile oraircraft in claim 9, wherein the closed-loop control system comprises aproportional-integral-derivative (PID) controller.
 11. The missile oraircraft in claim 9, wherein the closed-loop control system activatesand deactivates the at least one activatable flow effector or activeflow control device to create command forces on the afterbody tomaneuver the missile or aircraft.
 12. The missile or aircraft in claim8, wherein the at least one sensor is an inertial measurement unit. 13.The missile or aircraft in claim 7, further comprising a boattail andtail fins on the afterbody of the missile or aircraft, wherein the atleast one plasma actuator is located on the boattail or tail fins of themissile or aircraft.
 14. A missile or aircraft comprising: an afterbodyand a forebody; and at least one electromechanical activatable floweffector or electromechanical active flow control device on the missileor aircraft afterbody, and at least one sensor co-located with the atleast one activatable flow effector or active flow control device on themissile or aircraft afterbody; wherein the sensor is positioned todetect, and capable of detecting, flow separation from a surface of theafterbody and wherein the at least one activatable flow effector oractive flow control device is configured and located on the missile oraircraft afterbody to change a moment of the missile or aircraft, themoment being a pitching or yawing moment.
 15. The missile or aircraft ofin claim 14, further comprising a closed-loop control system, whereinthe closed-loop control system may be used for activating anddeactivating the at least one activatable flow effector or active flowcontrol device to change the moment.
 16. The missile or aircraft inclaim 15, wherein the closed-loop control system activates anddeactivates the at least one activatable flow effector or active flowcontrol device to create command forces on the afterbody to maneuver themissile or aircraft.
 17. The missile or aircraft in claim 14, whereinthe at least one activatable flow effector or active flow control deviceis a plasma actuator.
 18. The missile or aircraft in claim 14, whereinthe at least one activatable flow effector or active flow control deviceis a deployable flow effector.
 19. The missile or aircraft in claim 18,wherein the at least one activatable flow effector or active flowcontrol device, when deactivated, is flush, or nearly flush, with thesurface of the missile or aircraft.